Automated propeller feather testing

ABSTRACT

There is described herein the automation of propeller feather testing functions, whereby the test is automatically performed and a pass/fail signal is issued upon completion.

CROSS REFERENCE TO RELATED APPLICATIONS AND CLAIM OF PRIORITY

This application is a continuation of U.S. patent application Ser. No.16/655,424 filed on Oct. 17, 2019, which is a continuation of U.S.patent application Ser. No. 14/712,985 filed on May 15, 2015 and issuedas U.S. Pat. No. 10,487,682, the entire contents of which are herebyincorporated by reference.

TECHNICAL FIELD

The present invention relates to the field of propeller feathering andmore particularly, to dormancy tests for propeller feathering.

BACKGROUND OF THE ART

A feathered propeller has its blades moved to an extremely high pitchangle of approximately 90° so that they face perpendicular to theairstream and produce minimal aerodynamic drag. This may be doneintentionally during a flight to decrease the drag on an airplane and,prevent windmilling of the propeller. As this function is often used inemergency conditions in flight, regular testing of propeller featherfunctions is performed. Such testing is used to exercise the featheringmechanisms of the propeller, in order to ensure that there are nodormant failures present within the feather activation system. Thisactivation system may include any one of electronic, electrical,mechanical, and hydraulic features used to successfully feather thepropeller.

The feather test is conducted manually by a pilot, at engine start andtaxi-out of the aircraft. A push-button test switch is activated fromthe cockpit to command feathering of the propeller system. A successfulfeather test results in an audible drop in propeller speed which isdetectable by the pilot. The feather test switch is then released tocancel the feather test operation.

There is a need to improve propeller feather testing functions.

SUMMARY

There is described herein the automation of propeller feather testingfunctions, whereby the test is automatically performed and a pass/failsignal is issued upon completion. The automated propeller feather testmay be a system dormancy test and it may be performed while the aircraftis on the ground, during engine startup, shutdown, or other phases ofengine operation.

In accordance with a first broad aspect, there is provided a method fortesting a propeller feathering function. The method comprises monitoringa rotational speed over time of propeller blades of an aircraft;automatically commanding an angle change of the propeller blades;comparing a post-angle change rotational speed of the propeller bladesto an expected rotational speed without the commanded angle change andobtaining a rotational speed difference; and issuing a test passedsignal when the rotational speed difference exceeds a threshold and atest failed signal when the rotational speed difference does not exceedthe threshold.

In accordance with another broad aspect, there is provided a system fortesting a propeller feathering function, the system comprising: amemory; a processor coupled to the memory; and an application stored inthe memory. The application comprises program code executable by theprocessor for monitoring a rotational speed over time of propellerblades of an aircraft; automatically commanding an angle change of thepropeller blades; comparing a post-angle change rotational speed of thepropeller blades to an expected rotational speed without the commandedangle change and obtaining a rotational speed difference; and issuing atest passed signal when the rotational speed difference exceeds athreshold and a test failed signal when the rotational speed differencedoes not exceed the threshold.

In accordance with another broad aspect, there is provided a system fortesting a propeller feathering function of an aircraft. The systemcomprises a propeller; and a propeller control system. The propellercontrol system comprises an actuator coupled to the propeller forsetting a blade pitch of the propeller and a feathering test systemcoupled to the actuator and comprising a combination of software andhardware logic. The logic is configured for monitoring a rotationalspeed over time of the propeller; automatically commanding a change ofthe blade pitch; comparing a post-change rotational speed to an expectedrotational speed without the commanded change in blade pitch andobtaining a rotational speed difference; and issuing a test passedsignal when the rotational speed difference exceeds a threshold and atest failed signal when the rotational speed difference does not exceedthe threshold.

BRIEF DESCRIPTION OF THE DRAWINGS

Further features and advantages of the present invention will becomeapparent from the following detailed description, taken in combinationwith the appended drawings, in which:

FIG. 1 is a schematic side cross-sectional view of an exemplary gasturbine engine;

FIG. 2 is a block diagram of an exemplary aircraft propeller controlsystem;

FIG. 3 is a flowchart of an exemplary method for testing a propellerfeathering function;

FIG. 4 is a graph of propeller rotational speed versus time for variousembodiments of testing the propeller feathering function;

FIG. 5 is a block diagram of an exemplary embodiment for the featheringtest system; and

FIG. 6 is a block diagram of an exemplary embodiment of an applicationrunning on the processor of the feathering test system.

It will be noted that throughout the appended drawings, like featuresare identified by like reference numerals.

DETAILED DESCRIPTION

FIG. 1 illustrates an exemplary engine 10, namely a gas turbine engine,comprising an inlet 12, through which ambient air is propelled, acompressor section 14 for pressurizing the air, a combustor 16 in whichthe compressed air is mixed with fuel and ignited for generating anannular stream of hot combustion gases, and a turbine section 18 forextracting energy from the combustion gases. The turbine section 18illustratively comprises a compressor turbine 20, which drives thecompressor assembly and accessories, and at least one power or freeturbine 22, which is independent from the compressor turbine 20 anddrives the rotor shaft 24 through the reduction gearbox 26. Hot gasesmay then be evacuated through exhaust stubs 28. A rotor 30, in the formof a propeller through which ambient air is propelled, is hosted in apropeller hub 32. Rotor 30 may, for example, comprise a propeller of afixed-wing aircraft or a main (or tail) rotor.

The aircraft engine 10 may be used in combination with an aircraftpropeller control system 200, comprising an actuator 202 for modifyingblade pitch for propeller feathering, and a feathering test system 204,as illustrated in FIG. 2 . The propeller 30 converts rotary motion fromthe engine 10 to provide propulsive force to the aircraft. The pitch ofthe propeller 30 is variable and may be modified by the actuator 202.The actuator 202 may take different forms, depending on the type ofengine and/or aircraft. In some embodiments, there may also be gearing,such as that found on turboprop aircraft. The actuator 202 may rotatethe blades of the propeller 30 parallel to airflow in order to reducedrag in case of an engine failure. Such rotation may take the form of anincrease (towards feathered position) or a decrease (away from featheredposition) in blade pitch. The effect of an increase in blade pitch is toincrease the gliding distance of the aircraft and in some cases, tomaintain altitude with reduced engine power. The effect of a decrease inblade pitch is to help slow down an aircraft after landing in order tosave wear on the brakes and tires.

The feathering test system 204 is coupled to the actuator 202 andconfigured to perform automatic testing of a propeller featheringfunction of an aircraft, as illustrated in the exemplary method 300 ofFIG. 3 . As per step 302, an angle change of the propeller blades iscommanded automatically. The commanded angle change may be performed atany time prior to take-off and/or after landing, i.e. at any momentwhile the aircraft is on the ground. A trigger signal may be used toinitiate the commanded angle change. For example, the feathering testmay be associated with another task or procedure performed within theaircraft, such as engine shutdown, engine startup, or another procedureor test regularly performed by the aircraft in preparation for takeoffor after landing. Initiation of the associated task or procedure maygenerate the trigger signal and cause the automatic change in pitchangle for the propeller blades. The trigger signal may vary as afunction of the aircraft model, the engine type, the operatingenvironment, and internal policies/regulations of a given airline.

In some embodiments, step 302 may be preceded by detecting anaircraft-on-ground condition. Detection of an aircraft-on-groundcondition may be done using various techniques, such as aweight-on-wheels signal, a ground sensor, an airspeed sensor and aglobal positioning system. Other techniques may also be used. In suchcircumstances, step 302 may be performed conditionally upon detection ofthe aircraft-on-ground condition.

As per step 303, the rotational speed of the propeller blades ismonitored in order to detect a change subsequent to the commanded anglechange compared to an expected rotational speed. In particular, apost-angle change rotational speed of the propeller blades is comparedto the expected rotational speed of the propeller blades had the anglechange not occurred, and a rotational speed difference is obtained, asper step 304. The rotational speed difference may be compared to athreshold. A rotational speed difference that meets the threshold isindicative that the feathering function is operational. If therotational speed difference exceeds (or meets) the threshold, a testpass signal is issued, as per step 306. If the rotational speeddifference does not exceed (or does not meet) the threshold, a testfailed signal is issued, as per step 308.

Monitoring of the rotational speed may be performed using varioussensors, already present on the aircraft and used for other purposes, ordedicated to the automated feathering test. In some embodiments, themethod comprises returning the blade pitch to a zero pitch angle afterthe given time period.

In some embodiments, the blade pitch is moved from an initial zero pitchangle to a target pitch angle that is greater than a zero pitch angleand up to a maximum pitch angle (90°), such as but not limited to 5°,30°, 45°, and 70°. This is referred to as an increase in blade pitch asthe blades are moved towards the feathering position. In someembodiments, the blade pitch is moved from an initial zero pitch angleto a target pitch angle that is less than a zero pitch angle and up to aminimum pitch angle (−90°), such as but not limited to −5°, −30°, −45°,and −70°. This is referred to as a decrease in blade pitch as the bladesare moved away from the feathering position. In some embodiments, theblade pitch may be increased or decreased from a position other than azero pitch angle, and the rotational speed post-angle change is comparedto the expected rotational speed for the blades at the pre-angle changeposition.

In some embodiments, the blade pitch is set to the target pitch anglewith a single command. A timer may be used to set an end time for thetest. In other words, if a rotational speed difference greater than orequal to the threshold is not detected after a given time period, thetest is considered to have failed. The timer may be set for a givennumber of seconds, minutes, or any other unit of time as appropriate.

Alternatively, the blade pitch may be progressively changed until thetarget blade pitch is reached. The target blade pitch for the test mayvary as a function of the aircraft model, the engine type, the operatingenvironment, and internal policies/regulations of a given airline. Thetarget blade pitch may be fixed or may be programmable. Progressivechange of the blade pitch may be used in combination with the timer. Inother embodiments, the blade pitch may be progressively changed until itreaches maximum/minimum pitch or until the rotational speed differencemeets the threshold, whichever occurs first. The threshold may be set asdesired, such as a 5% change, a 10% change, a 25% change, or any otherappropriate amount.

An exemplary embodiment of performing the feathering test at engineshutdown is illustrated in FIG. 4 . Curve 402 represents the propellerrotational speed over time at zero pitch angle before engine shutdown.Curve 404 represents the expected propeller rotational speed over timeat zero pitch angle after engine shutdown. A natural decay rate for thespeed occurs due to the removal of rotary motion from the engine. Curve406 represents the post-angle change rotational speed over time atmaximum pitch angle, i.e. when the blade pitch has been set to 90°,after engine shutdown. Compared to curve 404, the decay rate for curve406 is shown to be significantly greater, as the increased rotationaldrag of the propeller caused by the change in blade pitch results in afaster decrease in rotational speed of the propeller. Curve 408represents the post-angle change rotational speed over time at a pitchangle that is greater than 0° and smaller than 90°, after engineshutdown. As illustrated, a slight change in pitch angle may besufficient to detect a rotational speed difference that meets athreshold in order to confirm that the propeller feathering function isoperational.

In some embodiments, monitoring the rotational speed over time comprisesmonitoring a rate of change of the rotational speed over time. Inaddition, comparing a post-angle change rotational speed of thepropeller blades to an expected rotational speed comprises comparing therate of change of the rotational speed of the propeller blades to anexpected rate of change of a zero pitch angle propeller. If performed atengine shutdown, this may comprise comparing the decay rate of therotational speed of the propeller blades at the target blade pitch to anexpected natural decay rate of a zero pitch angle propeller, as percurve 404. Coordinating the feathering test with engine shutdown allowsa common baseline to be used for same aircraft, as the natural decayrate may be consistent between the aircraft. Alternatively, thecommanded angle change may be triggered with delay from engine shutdown.This allows the decay rate of the rotational speed (or the rotationalspeed itself) before and after the commanded angle change to becompared, for detection of a change indicative of a successfulfeathering test. This embodiment is illustrated in FIG. 4 with curve410.

In some embodiments, the feathering test system 204 comprises acombination of hardware and software logic for performing the automatictesting. The feathering test system 204 may be a stand-alone unit or itmay be incorporated into existing aircraft systems architecture. Forexample, the feathering test system 204 may be part of an engine controlsystem, such as an electronic engine control (EEC) or a full authoritydigital engine control (FADEC). It may also be part of an integratedelectronic engine and propeller control system. In some embodiments, thefeathering test system 204 comprises a microcontroller and memory. Thememory may be SRAM, EEPROM, or Flash and the system 204 may be analogand/or digital based. Sensor values may be read by the microcontrollerand data may be interpreted using one or more lookup table. In someembodiments, the feathering test system 204 is programmable andcomprises a microprocessor which can process the inputs from enginesensors in real-time. Hardware may comprise electronic components on aprinted circuit board (PCB), ceramic substrate, or thin laminatesubstrate, with a microcontroller chip as a main component. Softwarecode may be stored in the microcontroller or other chips, and may beupdated by uploading new code or replacing the chips.

FIG. 5 illustrates another exemplary embodiment for the feathering testsystem 204. The feathering test system 204 may comprise, amongst otherthings, a plurality of applications 506 a . . . 506 n running on aprocessor 504 coupled to a memory 502. It should be understood thatwhile the applications 506 a . . . 506 n presented herein areillustrated and described as separate entities, they may be combined orseparated in a variety of ways. The memory 502 accessible by theprocessor 504 may receive and store data. The memory 502 may be a mainmemory, such as a high speed Random Access Memory (RAM), or an auxiliarystorage unit, such as a hard disk, a floppy disk, or a magnetic tapedrive. The memory 502 may be any other type of memory, such as aRead-Only Memory (ROM), or optical storage media such as a videodisc anda compact disc. The processor 504 may access the memory 502 to retrievedata. The processor 504 may be any device that can perform operations ondata. Examples are a central processing unit (CPU), a front-endprocessor, a microprocessor, and a network processor. The applications506 a . . . 506 n are coupled to the processor 304 and configured toperform various tasks.

FIG. 6 is an exemplary embodiment of an application 506 a running on theprocessor 504. The application 506 a illustratively comprises acommanding module 602, a monitoring module 604, and a test result module606. The commanding module 602 is configured for receiving the triggersignal and automatically commanding the angle change of the propellerblades, via the actuator 202. Once the blade pitch has been modified bythe commanding module, the monitoring module 604 may be advised by thecommanding module 602. Alternatively, the monitoring module 604 may beconfigured to continuously monitor rotational speed of the propellerblades, or it may itself receive the trigger signal and begin monitoringbefore the commanding module 602 is instructed to command the anglechange. The monitoring module 604 receives from various sensors inputdata for comparing a post-angle change rotational speed of the propellerblades to an expected rotational speed without the commanded anglechange and obtaining a rotational speed difference. The monitoringmodule 604 is connected to the test result module 606 for transmittingthe rotational speed difference thereto. If the rotational speeddifference meets the threshold, the test result module 606 will issue apass signal. If the rotational speed difference does not meet thethreshold, the test result module 606 will issue a fail signal.

In some embodiments, the test result module 606 is also configured toissue a maintenance required signal in case of a failed feathering test.The maintenance required signal may be generic and applicable to anyfailed feathering test. Alternatively, different maintenance requiredsignals may be provided as a function of the specifics of the featheringtest. For example, the monitoring module 604 may be configured todetermine if the problem is related to the electronics, the actuator,oil, or the propeller blades themselves. This information may be passedon to the test result module 606 and the appropriate maintenancerequired signal may be issued accordingly. The pass/fail signal mayresult in a visual indicator for the pilot and/or the ground crew, suchas a red light for a failed test and a green light for a passed test. Amaintenance required signal may be part of the same visual indicator asa failed test or may result in a separate visual indicator for the pilotand/or ground crew.

Other variants to the configurations of the commanding module 602, themonitoring module 604, and the test result module 606 may also beprovided and the example illustrated is simply for illustrativepurposes.

The above description is meant to be exemplary only, and one skilled inthe relevant arts will recognize that changes may be made to theembodiments described without departing from the scope of the inventiondisclosed. For example, the blocks and/or operations in the flowchartsand drawings described herein are for purposes of example only. Theremay be many variations to these blocks and/or operations withoutdeparting from the teachings of the present disclosure. For instance,the blocks may be performed in a differing order, or blocks may beadded, deleted, or modified. While illustrated in the block diagrams asgroups of discrete components communicating with each other via distinctdata signal connections, it will be understood by those skilled in theart that the present embodiments are provided by a combination ofhardware and software components, with some components being implementedby a given function or operation of a hardware or software system, andmany of the data paths illustrated being implemented by datacommunication within a computer application or operating system. Thestructure illustrated is thus provided for efficiency of teaching thepresent embodiment. The present disclosure may be embodied in otherspecific forms without departing from the subject matter of the claims.Also, one skilled in the relevant arts will appreciate that while thesystems, methods and computer readable mediums disclosed and shownherein may comprise a specific number of elements/components, thesystems, methods and computer readable mediums may be modified toinclude additional or fewer of such elements/components. The presentdisclosure is also intended to cover and embrace all suitable changes intechnology. Modifications which fall within the scope of the presentinvention will be apparent to those skilled in the art, in light of areview of this disclosure, and such modifications are intended to fallwithin the appended claims.

The invention claimed is:
 1. A method for testing a propeller pitchangle function, the method comprising: while propeller blades of anaircraft are rotating about a rotational axis, commanding, by aprocessor of a computing device, a pitch angle change of the propellerblades; after said pitch angle change of the propeller blades, measuringa post-angle change rotational speed of the propeller blades; comparing,by the processor of the computing device, the post-angle changerotational speed of the propeller blades to an expected rotational speedwithout the commanded angle change; and issuing, by the processor of thecomputing device and based on said comparing, one of a test passedsignal indicative of a passed pitch angle test and a test failed signalindicative of a failed pitch angle test.
 2. The method of claim 1,wherein said commanding is performed following a change in an enginestate.
 3. The method of claim 2, wherein the change of the engine stateis one of an initiation of an engine shutdown and an initiation of anengine start-up.
 4. The method of claim 1, further comprising detectingan aircraft-on-ground condition, and wherein said commanding the pitchangle change of the propeller blades comprises commanding the pitchangle change only when the aircraft-on-ground condition is detected. 5.The method of claim 1, further comprising returning the propeller bladesto a zero pitch angle after issuing the test passed signal.
 6. Themethod of claim 1, wherein said commanding the pitch angle change of thepropeller blades comprises commanding the pitch angle change to a targetblade pitch that is less than a maximum blade pitch.
 7. The method ofclaim 1, wherein said commanding the pitch angle change of the propellerblades comprises progressively changing a pitch angle of the propellerblades until a target blade pitch is reached.
 8. The method of claim 7,wherein comparing the post-angle change rotational speed of thepropeller blades to an expected rotational speed comprises comparinguntil reaching the target blade pitch.
 9. The method of claim 1, whereinsaid commanding the pitch angle change of the propeller blades comprisesincreasing a blade pitch, and wherein comparing the post-angle changerotational speed to the expected rotational speed comprises detecting adecrease in post-angle change rotational speed compared to the expectedrotational speed.
 10. The method of claim 1, wherein said commanding thepitch angle change of the propeller blades comprises decreasing a bladepitch, and wherein comparing the post-angle change rotational speed tothe expected rotational speed comprises detecting an increase inpost-angle change rotational speed compared to the expected rotationalspeed.
 11. The method of claim 1 wherein said comparing comprisesobtaining a rotational speed difference, and wherein said issuingcomprises issuing a test passed signal when the rotational speeddifference exceeds a threshold and a test failed signal when therotational speed difference does not exceed the threshold.
 12. A systemfor testing a propeller pitch angle function, the system comprising: amemory; a processor coupled to the memory; and an application stored inthe memory and comprising program code executable by the processor for:while propeller blades of an aircraft are rotating about a rotationalaxis, commanding a pitch angle change of the propeller blades; aftersaid pitch angle change of the propeller blades, measuring a post-anglechange rotational speed of the propeller blades; comparing thepost-angle change rotational speed of the propeller blades to anexpected rotational speed without the commanded angle change; andissuing, based on said comparing, one of a test passed signal indicativeof a passed pitch angle test and a test failed signal indicative of afailed pitch angle test.
 13. The system of claim 12, wherein saidcommanding is performed following a change in an engine state, thechange in the engine state being one of an initiation of an engineshutdown and an initiation of an engine start-up.
 14. The system ofclaim 12, further comprising detecting an aircraft-on-ground condition,and wherein said commanding the pitch angle change of the propellerblades comprises commanding the pitch angle change only when theaircraft-on-ground condition is detected.
 15. The system of claim 12,further comprising returning the propeller blades to a zero pitch angleafter issuing the test passed signal.
 16. The system of claim 12,wherein said commanding the pitch angle change of the propeller bladescomprises commanding the pitch angle change to a target blade pitch thatis less than a maximum blade pitch.
 17. The system of claim 12, whereinsaid commanding the pitch angle change of the propeller blades comprisesprogressively changing a pitch angle of the propeller blades until atarget blade pitch is reached.
 18. A method for testing a propellerpitch angle function, the method comprising: while propeller blades ofan aircraft are rotating at a rotational speed about a rotational axis,commanding, by a processor of a computing device, a pitch angle changeof the propeller blades; after said pitch angle change of the propellerblades, measuring a post-angle rate of change of the rotational speed ofthe propeller blades; comparing, by the processor of the computingdevice, the post-angle change rate of change of the rotational speed toan expected rate of change of the rotational speed without the commandedpitch angle change; and issuing, by the processor of the computingdevice and based on said comparing, one of a test passed signalindicative of a passed pitch angle test and a test failed signalindicative of a failed pitch angle test.
 19. The method of claim 18,wherein said commanding is performed following a change in an enginestate, the change of the engine state being an initiation of an engineshutdown.
 20. The method of claim 18, wherein said commanding isperformed following a change in an engine state, the change of theengine state being an initiation of an engine start-up.